The present invention relates to tooling designed to repair a composite structure with a damaged zone without the structure having to be displaced to a repair site different from that in which the structure is located.
The term xe2x80x9ccomposite structurexe2x80x9d as used here and in the text that follows refers to a structure composed of long fibers, such as carbon or other fibers, arranged in superimposed layers or folds that are embedded in a resin matrix.
The invention also relates to a method for repairing a composite structure using this kind of tooling.
The tooling and method of the invention may be used in any situation where it is not possible or too impractical to send the structure to a specialized workshop or put it through an autoclave. One of the principle applications for the tooling and method of the invention thus relates to repairing composite structures on aircraft.
Composite materials enable mechanical characteristics to be obtained that are perfectly controlled and comparable to those obtained using metal, while being considerably lighter. Furthermore, the manufacturing techniques of these materials are continually being improved, both in terms of the shapes and measurements of the parts and the length of manufacture time and cost. It is for these various reasons that composite structures are being used more and more frequently in aircraft.
However, compared to metal structures, composite structures have the drawback of being more fragile when they are subjected to impact. When a composite structure is violently hit by an object variable degrees of breaking or cracking occur in the zone of impact. The increase in the number of composite structures used in aircraft has therefore led to significant improvements in the techniques used to repair these structures.
Two criteria should be adopted when the quality of a repair on an aircraft composite structure is evaluated: the appearance of the repaired zone and the mechanical characteristics of the structure in this zone. Furthermore, it is essential for the repair work to be carried out on-site, i.e. in the place where the aircraft is located, in order for costs to be reduced and to limit aircraft downtime to a minimum. In practice this rules out the use of autoclaves.
Appearance is an important criterion given that the damaged zone is usually located on the outside of the aircraft, i.e. in a zone that can be seen. It is therefore preferable for the zone to resemble the composite structure before it was damaged, such that it is practically invisible.
Moreover, the larger the damaged zone, the more crucial it becomes for the mechanical characteristics of the repaired zone to be as close as possible to those initially present in the composite structure before it was damaged.
At present a number of devices exist that enable composite structure repairs to be carried out on-site when the damaged zone is relatively small. These devices include tooling for step machining or scarf machining a recess in the damaged zone of the composite structure and tooling used to polymerize a composite part placed in the recess under pressure.
U.S. Pat. Nos. 5,207,541 and 5,271,145 concern machining tooling. Tooling of this kind is generally used to machine circular- or oval-shaped recesses. However, other shapes are also possible, as disclosed by U.S. Pat Nos. 4,916,880 and 4,978,404.
The composite part set into a recess as described is of a shape that matches that of the recess. The composite part generally comprises several layers or folds of long fibre fabric and non-polymerized resin. If appropriate, the fabric may previously be impregnated with resin or be initially dry. If the fabric is dry the resin can either be applied as a liquid on each layer of fabric or as solid film interposed between the various layers of dry fabric.
Generally, the tooling used to polymerize and compact the composite part mainly comprises a heating cap that raises the resin to the temperature at which it polymerizes and a bladder fastened around the composite part such that it forms a leaktight seal on the surface of the composite structure. The bladder is connected to an external source of negative pressure such that pressure may be applied to the composite part.
As shown in U.S. Pat. No. 4,554,036, the pressure applied to the composite part in this standard tooling remains very reduced, for example approximately 1.5 bar.
This kind of tooling can therefore be used for xe2x80x9ccosmeticxe2x80x9d repairs, in other words for reconstituting the original appearance of the composite structure. However, this tooling affects the mechanical characteristics of the repaired zone and can reduce the characteristics by up to 30% compared to the structure in its original condition. Repair work using this technique is, therefore, limited to only small damaged zones.
When larger zones are damaged the damaged composite structure is usually completely replaced and then sent to be repaired in a specialist workshop equipped with autoclaves that perform high pressure polymerization under appropriate conditions.
In another known technique a composite part is molded in an autoclave into a shape that matches that of the recess previously machined in the damaged zone. The composite part is then polymerized and compacted in an autoclave before being bonded into place in the recess. This technique is discussed in American patent 5 023 87.
The major drawback of this technique is that the ability to withstand fatigue in the bonded zone is relatively limited, thus constituting a particular drawback for the aeronautics industry. It is for this reason that the fastening of the additional composite part is generally completed by other fastening means, such as bolts or rivets, that require access via the rear surface of the structure. The repair work can be seen from the outside and does not, therefore, meet the above-mentioned requirements concerning appearance.
Furthermore, the polymerized composite part is manufactured in an autoclave, i.e. generally on a different site from that where the aircraft is located. This means that the aircraft will be immobilized for a long and costly period.
The invention mainly relates to tooling for on-site repairs of a composite structure with a damaged zone. The tooling restores the original appearance of the composite structure and guarantees mechanical characteristics that are very similar to those of the undamaged structure, thus allowing on-site repairs of relatively large damaged zones.
According to the invention, these results are obtained using on-site repair tooling for composite structures with a concave recess previously machined in a surface of the structure for high-pressure polymerization of a non-polymerized composite part. The shape of the composite part matches that of the recess and the tooling is characterized in that it comprises:
a bladder capable of being fastened around the said composite part such that it forms a leaktight seal on the surface of the composite structure;
means for connecting the bladder to an external source of negative pressure;
means for heating the composite part; and
means for increasing the pressure applied to the composite part, said means being suitable for being interposed between said composite part and the leaktight bladder.
In a preferred embodiment of the invention the means for increasing pressure comprise a stack of at least two plates with surfaces that gradually increase towards the bladder.
For example, the surfaces of the plates the furthest away from the stack are in a ratio of at least 1:2.
To take into account the fact that that the surface of an aircraft composite structure is not generally plane the plates are preferably made of plane leaf metal that can be stretched to take the shape of the composite structure.
Guiding means are provided to hold the plates in a centered position in relation to each other when pressure is applied by creating a vacuum in the bladder.
These guiding means may comprise at least one guiding pin that projects perpendicularly out of at least one of the plates through guiding holes provided in the other plates.
The surface and the shape of the plate that is closest to the composite part are preferably more or less identical to those of said composite part on the surface of the structure.
According to a characteristic known in the art, the heating means comprise a heating cap that is interposed, in this example, between the bladder and the means for increasing pressure.
The invention also relates to a method for on-site repairs of a composite structure with a damaged zone, characterized in that it comprises the following stages:
machining a concave recess in the damaged zone in a surface of the structure;
setting a non-polymerized composite part in the recess, the shape of said composite part matching that of the recess;
polymerizing the composite part in situ using a vacuum created inside a bladder that is sealed onto the said surface, around said composite part, after means for increasing the pressure applied to the composite part have been interposed between the part and the bladder; heating of the composite part.